Method and system to facilitate enhanced local cooling of turbine engines

ABSTRACT

A method of assembling a gas turbine engine is provided. The method includes coupling at least one turbine nozzle segment within the gas turbine engine. The at least one turbine nozzle segment includes at least one airfoil vane extending between an inner band and an outer band that includes an aft flange and a radial inner surface. The method also includes coupling at least one turbine shroud segment downstream from the at least one turbine nozzle segment, wherein the at least one turbine shroud segment includes a leading edge and a radial inner surface, coupling a cooling fluid source in flow communication with the at least one turbine nozzle segment such that cooling fluid channeled to each turbine nozzle outer band aft flange is directed at an oblique discharge angle towards the leading edge of the at least one turbine shroud segment, and channeling the cooling fluid through at least a first group of cooling openings having a larger aggregate cross-sectional area and a second group of cooling openings having a smaller aggregate cross-sectional area to facilitate preferential cooling of the turbine shroud.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH & DEVELOPMENT

The U.S. Government may have certain rights in this invention pursuantto contract number N00019-04-C-0093.

BACKGROUND OF THE INVENTION

This invention relates generally to gas turbine engines and, moreparticularly, to methods and systems for cooling integral turbine nozzleand shroud assemblies.

One known approach to increase the efficiency of gas turbine enginesrequires raising the turbine operating temperature. However, asoperating temperatures are increased, the thermal limits of certainengine components may be exceeded, resulting in reduced service lifeand/or material failure. Moreover, the increased thermal expansion andcontraction of components may adversely affect component clearancesand/or component interfitting relationships. Consequently, conventionalcooling systems have been incorporated into gas turbine engines tofacilitate cooling such components to avoid potentially damagingconsequences when exposed to elevated operating temperatures.

It is known to extract, from the main airstream, air from the compressorfor cooling purposes. To facilitate maintaining engine operatingefficiency, the volume of cooling air extracted is typically limited toa small percentage of the total main airstream. As such, this requiresthat the cooling air be utilized with the utmost efficiency in order tofacilitate maintaining the temperatures of components within safelimits.

For example, one component that is subjected to high temperatures is theshroud assembly located immediately downstream of the high pressureturbine nozzle extending from the combustor. The shroud assembly extendscircumferentially about the rotor of the high pressure turbine and thusdefines a portion of the outer boundary (flow path) of the main gasstream flowing through the high pressure turbine. Gas turbine engineefficiency may be negatively affected by a fluctuation in turbine bladetip clearance measured between a radially outer surface of the turbineblade and a radially inner surface of the shroud assembly. Duringtransient engine operation, turbine blade tip clearance is a function ofa difference in radial displacement of the turbine rotor blade and theshroud assembly. The turbine rotor typically has a larger mass than thestationary shroud system and, thus, during turbine operation, theturbine rotor typically has a slower thermal response than the shroudassembly. When the difference in the rotor blade radial displacement andthe shroud assembly radial displacement is too great, the blade tipclearance is increased, which may result in reducing engine efficiency.

Moreoever, during engine operation, a gap may be defined between atrailing edge of the high pressure turbine nozzle outer band and aleading edge of the adjacent shroud segment. Cooling air, including,without limitation, nozzle leakage and/or purge flow, enters the gap andflows into the main gas stream channeled through the high pressureturbine. Cooling air is generally provided by a row of axially alignedcooling holes positioned in an outer band trailing edge that aredirected towards the shroud leading edge forward face to facilitatecooling the end faces and purging the gap. Because known nozzle outerband trailing edges and shroud leading edges have a simple 90° corner,the gap opens directly into the main gas stream. During engineoperation, as the main gas stream flows through the nozzle vanes, acircumferential gas pressure variation may be created downstream fromthe vane trailing edge. This circumferential gas pressure variation maycause localized hot gas ingestion into the gap between the outer bandand the shroud segment. As a result, cooling air flowing through the gapmay not effectively cool the downstream shroud segement.

BRIEF DESCRIPTION OF THE INVENTION

In one aspect, a method of assembling a gas turbine engine is provided.The method includes coupling at least one turbine nozzle segment withinthe gas turbine engine. The at least one turbine nozzle segment includesat least one airfoil vane extending between an inner band and an outerband that includes an aft flange and a radial inner surface. The methodalso includes coupling at least one turbine shroud segment downstreamfrom the at least one turbine nozzle segment, wherein the at least oneturbine shroud segment includes a leading edge and a radial innersurface, coupling a cooling fluid source in flow communication with theat least one turbine nozzle segment such that cooling fluid channeled toeach turbine nozzle outer band aft flange is directed at an obliquedischarge angle towards the leading edge of the at least one turbineshroud segment, and channeling the cooling fluid through at least afirst group of cooling openings having a larger aggregatecross-sectional area and a second group of cooling openings having asmaller aggregate cross-sectional area to facilitate preferentialcooling of the turbine shroud.

In another aspect, an engine assembly is provided. The engine assemblyincludes a nozzle assembly including an inner band, an outer bandincluding an aft flange and a radial inner surface. The aft flangeincludes a first group of cooling openings having a larger aggregatecross-sectional area and a second group of cooling openings having asmaller aggregate cross-sectional area. The cooling openings areconfigured to direct cooling fluid therefrom at an oblique dischargeangle. The assembly also includes at least one airfoil vane extendingbetween the inner band and the outer band.

In yet another aspect, a gas turbine engine is disclosed. The gasturbine engine includes a nozzle assembly including an inner band, anouter band, and at least one airfoil vane extending between the innerband and the outer band. The outer band includes an aft flange and aradial inner surface. The aft flange includes a first group of coolingopenings having a larger aggregate cross-sectional area and a secondgroup of cooling openings having a smaller aggregate cross-sectionalarea. The cooling openings are configured to direct cooling fluidtherefrom at an oblique discharge angle.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a side view of an exemplary shroud assembly schematicallyillustrating high pressure cooling air flow through the shroud assembly;

FIG. 2 is a side view of an alternative shroud assembly schematicallyillustrating high pressure cooling air flow through the shroud assembly;

FIG. 3 is an enlarged schematic cross-sectional view of a gap definedbetween a turbine nozzle and the shroud assembly shown in FIG. 1 or 2;

FIG. 4 is a plan view of the turbine nozzle and shroud assembly shown inFIG. 3, illustrating an exemplary cooling hole pattern;

FIG. 5 is a plan view of the turbine nozzle and shroud assembly shown inFIG. 3, illustrating an alternative exemplary cooling hole pattern;

FIG. 6 is a plan view of the turbine nozzle and shroud assembly shown inFIG. 3, illustrating another alternative exemplary cooling hole pattern;and

FIG. 7 is a plan view of the turbine nozzle and shroud assembly shown inFIG. 3, illustrating yet another alternative exemplary cooling holepattern.

DETAILED DESCRIPTION OF THE INVENTION

The present invention provides a turbine shroud cooling system forminimizing hot gas ingestion into a gap defined between a trailing edgeof the high pressure turbine nozzle and a leading edge of the adjacentshroud segment. The turbine shroud cooling system facilitates forming abarrier between the hot gas flow path flowing through the high pressureturbine and cooling air flowing through a gap defined between theturbine nozzle and the shroud segment.

Although the present invention is described below in reference to itsapplication in connection with cooling a shroud assembly of an aircraftgas turbine, it should be apparent to those skilled in the art andguided by the teachings herein provided that with appropriatemodification, the cooling system or assembly of the present inventioncan also be suitable to facilitate cooling other turbine enginecomponents, such as, but not limited to, the nozzle and/or vanesections.

FIG. 1 is a side view of an exemplary shroud assembly schematicallyillustrating high pressure cooling air flow through the shroud assembly.FIG. 2 is a side view of an alternative shroud assembly schematicallyillustrating high pressure cooling air flow through the shroud assembly.To facilitate controlling shroud assembly thermal response and/or shroudassembly displacement during transient engine operation, in theexemplary embodiment, a turbine engine cooling assembly 108 includes ashroud assembly, generally indicated as 110, for a high pressure turbinesection 112 and a low pressure turbine section 114 of a gas turbineengine. It should be apparent to those skilled in the art and guided bythe teachings herein provided that turbine engine cooling assembly 108may be suitable to facilitate cooling other sections of the gas turbineengine, such as, but not limited to, a nozzle section and/or a vanesection.

Shroud assembly 110 includes turbine engine cooling components in theform of shroud segments 130. Each shroud segment 130 includes a forwardmounting hook 132 at a circumferential leading edge 133 of shroudsegment 130. Shroud segment 130 also includes a midsection mounting hook134 and an aft mounting hook 136 adjacent to a circumferential trailingedge 137 of shroud segment 130.

A plurality of shroud segments 130 are arranged circumferentially in agenerally known fashion to form an annular segmented shroud. Shroudsegments 130 define an annular clearance between high pressure turbineblades (not shown) and a radially inner surface 138 of a high pressureturbine section of shroud segments 130, and between low pressure turbineblades (not shown) and a radially inner surface 140 of a low pressureturbine section of shroud segment 130. A plurality of segmented shroudsupports 144 interconnect shroud segments 130. Each shroud support 144circumferentially spans and supports adjacent shroud segments 130. Inalternative embodiments, shroud supports 144 are modified to support anysuitable number of shroud segments 130 less than or greater than twoshroud segments 130. In the exemplary embodiment, shroud assembly 110includes twenty-six (26) shroud segments 130 and thirteen (13) shroudsupports 144, although any suitable number of shroud segments 130 and/orshroud supports 144 may be utilized in alternative embodiments.

Each shroud support 144 includes a forward section 146, a midsection 148and an aft section 150 that form respective forwardly projecting hangers152, 154 and 156. Mounting hooks 132, 134 and 136 are received bycooperating hangers 152, 154 and 156, respectively, in tongue-in-groove,or hook-in-hanger, interconnections such that shroud support 144supports respective shroud segments 130.

Shroud assembly 110 includes an annular shroud ring structure 158 thatsupports shroud supports 144. In one embodiment, shroud ring structure158 is a one-piece, continuous annular shroud ring structure. A radialposition of each shroud support 144, as well as of each shroud segment130, is closely controlled by only two annular position control rings162 and 164 formed on shroud ring structure 158. In contrast toconventional shroud ring structures, to facilitate reducing or limitinga weight of shroud assembly 110, shroud ring structure 158 includes onlytwo position control rings 162 and 164. A midsection position controlring 162 includes an axially forwardly projecting hanger 166 thatreceives and/or cooperates with a rearwardly projecting mounting hook167 formed by support structure midsection 148 in a firstcircumferential tongue-in-groove or hook-in-hanger interconnection. Anaft position control ring 164 includes an axially forwardly projectinghanger 168 that receives and/or cooperates with a rearwardly projectingmounting hook 169 of support structure aft section 150 in secondcircumferential tongue-in-groove or hook-in-hanger interconnection.

In the exemplary embodiment, hangers 166 and/or 168 are in direct axialalignment, i.e., aligned generally in the same radial plane, withrespective hanger 154 and hanger 156 to facilitate maximizing the radialsupport and/or radial position control provided to shroud support 144and, thus, corresponding shroud segments 130. This alignment orientationfacilitates increasing the rigidity of the entire shroud supportassembly. In an alternative embodiment, shown in FIG. 2, hanger 166and/or hanger 168 are in an offset axial alignment, i.e., not alignedgenerally in the same radial plane, with respective hanger 154 andhanger 156. In the exemplary embodiment, shroud ring structure 158 isbolted to the combustor case (not shown) at an aft end of shroud ringstructure 158. Shroud ring structure 158 is cantilevered away fromleading edge 133 at the combustor case interface. As such, midsectionposition control ring 162 is positioned several inches away from thecombustor aft flange (not shown), and is thereby divorced from anynon-uniform circumferential variations in radial deflection in thecombustor case.

In the exemplary embodiment, high pressure cooling air 170 is extractedfrom a compressor (not shown) positioned upstream of shroud assembly110. A first portion 171 of high pressure cooling air 170 extracted fromthe compressor facilitates cooling high pressure turbine section 112. Asecond portion 172 of high pressure cooling air 170 extracted from thecompressor facilitates cooling low pressure turbine section 114.Referring further to FIG. 1, directional arrows corresponding to firstportion 171 and second portion 172 illustrate at least a portion of aflow path of first portion 171 of high pressure cooling air 170 througha high pressure turbine section active convection cooling zone 173 andsecond portion 172 of high pressure cooling air 170 through a lowpressure turbine section active convection cooling zone 186 (describedbelow), respectively.

In this embodiment, first portion 171 of high pressure cooling air 170is metered into a first or high pressure turbine section activeconvection cooling zone 173. More specifically, first portion 171 ofhigh pressure cooling air 170 is metered through at least one highpressure turbine section (HPTS) feed hole 174 defined in shroud support144. First portion 171 of high pressure cooling air 170 impinges againsta pan-shaped HPTS impingement baffle 175 positioned within high pressureturbine section active convection cooling zone 173. Baffle 175 iscoupled to shroud support 144 and thus at least partially defines anupper HPTS cavity or plenum 176. First portion 171 of high pressurecooling air 170 is then metered through a plurality of perforations 177formed in impingement baffle 175 as cooling air into a lower HPTS cavityor plenum 178 defined in shroud segment 130, wherein the cooling airimpinges against a backside 179 of shroud segment 130. A portion, suchas spent impingement cooling air 180, of high pressure cooling air exitsplenum 178 through a plurality of forwardly directed cooling openings181 defined at, or near, shroud segment leading edge 133 configured tofacilitate purging a gap 182 defined between high pressure turbinenozzle outer band 183 and shroud segment leading edge 133. A portion 184of high pressure cooling air is metered through a plurality ofrearwardly directed cooling openings 185 defined in shroud segment 130to facilitate film cooling inner surface 138 and/or 140. Spentimpingement cooling air 180 of high pressure cooling air exiting coolingopenings 181 facilitates preventing or limiting hot gas injection orrecirculation into shroud assembly 110 at leading edge 133.

Second portion 172 of high pressure cooling air 170 extracted from thecompressor facilitates cooling low pressure turbine section 114. In thisembodiment, second portion 172 of high pressure cooling air 170 ismetered into a second or low pressure turbine section active convectioncooling zone 186. More specifically, second portion 172 of high pressurecooling air 170 is metered through at least one low pressure turbinefeed hole 187 defined in shroud support 144. Second portion 172 of highpressure cooling air 170 impinges against a pan-shaped Low PressureTurbine Section impingement baffle 188 positioned within low pressureturbine section active convection cooling zone 186. Baffle 188 iscoupled to shroud support 144, and thus at least partially defines anupper LPTS cavity or plenum 189. Second portion 172 of high pressurecooling air 170 is then metered through perforations 190 defined inimpingement baffle 188 and into a lower LPTS cavity or plenum 191wherein high pressure cooling air impinges against a backside 192 ofshroud segment 130. Cooling air 193 exits plenum 191 through a pluralityof rearwardly directed cooling openings 194 defined through shroudsegment 130, to facilitate film cooling radially inner surface 140 oftrailing edge 137 of shroud segment 130 downstream.

As shown in FIG. 1, high pressure cooling air 170 is initially directedinto a duct 204 defined at least partially between high pressure turbinenozzle outer band 183 and the portion of shroud ring structure 158forming midsection position control ring 162. High pressure cooling air170 is separated within duct 204 into first portion 171, and into secondportion 172, as high pressure cooling air 170 is directed through duct204. First portion 171 of high pressure cooling air 170 is meteredthrough HPTS feed holes 174 into active convection cooling zone 173 andinto plenum 178 to facilitate impingement cooling in high pressureturbine section 112. Spent impingement cooling air 180 exits shroudsegment 130 through shroud segment leading edge cooling openings 181 tofacilitate purging gap 182 defined between high pressure turbine nozzleouter band 183 and shroud segment 130, and/or through cooling openings185 defined at a trailing end 205 of high pressure turbine section 112to facilitate film cooling inner surface 138 and/or 140 of shroudsegment 130.

Second portion 172 of high pressure cooling air 170 is directed intosecond active convection cooling zone 186 that is defined at leastpartially between shroud support 144 and shroud segment 130, and betweenmidsection position control ring 162 and aft position control ring 164.Second portion 172 of high pressure cooling air 170 facilitates coolinglow pressure turbine section 114. In one embodiment, second portion 172of high pressure cooling air 170 is metered through a plurality of lowpressure turbine feed holes 187 defined in shroud support 144. Morespecifically, second portion 172 of high pressure cooling air 170 ismetered directly into active convection cooling zone 186 to facilitateshroud segment impingement cooling in low pressure turbine section 114,such that cooling air bypasses a third region 210 defining an inactiveconvection cooling zone 211 between shroud support 144 and shroud ringstructure 158, and between midsection position control ring 162 and aftposition control ring 164. Spent impingement cooling air exits shroudsegment 130 through cooling openings 194 defined at or near trailingedge 137 of shroud segment 130.

In the flow path illustrated in FIG. 1, high pressure turbine sectionactive convection cooling zone 173 and/or low pressure turbine sectionactive convection cooling zone 186 are directly and actively cooled. Lowpressure turbine section inactive convection cooling zone 211 isinactive, i.e., no high pressure cooling air flows through inactiveconvection cooling zone 211. Thus, a thermal response within inactiveconvection cooling zone 211 to environmental conditions created duringtransient engine operation is reduced and/or retarded. As a result,transient displacement of midsection position control ring 162 and/oraft position control ring 164 is also reduced and/or retarded.

As shown in FIG. 2, high pressure cooling air 170 is directed into duct204 defined at least partially between high pressure turbine nozzleouter band 183 and shroud ring structure 158 forming midsection positioncontrol ring 162. High pressure cooling air 170 is separated into firstportion 171 and second portion 172. First portion 171 of high pressurecooling air 170 is metered through HPTS feed hole(s) 174 into highpressure turbine section active convection cooling zone 173 at leastpartially defining plenum 176 and plenum 178 to facilitate shroudsegment impingement cooling in high pressure turbine section 112. Spentimpingement cooling air 180 exits shroud segment 130 through shroudsegment leading edge cooling openings 181 to facilitate purging gap 182between high pressure turbine nozzle outer band 183 and shroud segment130 and/or through cooling openings 185 defined at trailing end 205 ofhigh pressure turbine section 112 to facilitate film cooling innersurface 138 and/or 140.

Second portion 172 of high pressure cooling air 170 is directed into lowpressure turbine section active convection cooling zone 186 defined atleast partially between shroud support 144 and shroud segment 130, andbetween midsection position control ring 162 and aft position controlring 164 to facilitate cooling low pressure turbine section 114. In oneembodiment, second portion 172 of high pressure cooling air 170 ismetered through a plurality of low pressure turbine feed holes 187defined through shroud support 144. Second portion 172 of high pressurecooling air 170 is metered directly into low pressure turbine sectionactive convection cooling zone 186 at least partially defining plenum189 and plenum 191 to facilitate shroud segment impingement cooling inlow pressure turbine section 114. Spent impingement cooling air 193exits shroud segment 130 through cooling openings 194 defined at or neartrailing edge 137 of shroud segment 130.

The shroud cooling assembly as shown in FIGS. 1 and 2 directs highpressure cooling air directly into high pressure turbine section activeconvection cooling zone 173 and/or low pressure turbine section activeconvection cooling zone 186 through respective feed hole(s) 174 and feedhole(s) 187.

In the shroud cooling assembly as shown in FIGS. 1 and 2, high pressurecooling air is not metered or directed through low pressure turbinesection inactive convection cooling zone 211. As a result, thecomponents defining low pressure turbine section inactive convectioncooling zone 211 respond relatively slower to thermal conditions and/orenvironments during transient engine operation than the componentsdefining an active convection cooling zone within conventional shroudcooling assemblies. This slower response to thermal conditions and/orenvironments facilitates relatively slower transient displacement ofmidsection position control ring 162 and/or aft position control ring164.

Thus, by bypassing the low pressure turbine section shroud ringstructure, the high pressure cooling air flow paths shown in FIGS. 1 and2 facilitate reducing and/or retarding the transient thermal responseand/or displacement of the shroud segment during transient engineoperation. The slower response further facilitates improved blade tipclearance and turbine engine efficiency.

FIG. 3 is an enlarged schematic cross-sectional view of turbine nozzleband 183, gap 182 and shroud segment leading edge 133. Turbine nozzleouter band 183 is included as part of a turbine nozzle segment 520.Turbine nozzle segments 520 generally include a plurality ofcircumferentially-spaced airfoil vanes 510, shown in FIG. 4. The vanes510 extend between radial outer band 183 and a radial inner band (notshown). In the exemplary embodiment, outer band 183 includes a radialinner surface 522 and an aft flange 504. Aft flange 504 includes anupstream face 506, trailing edge 500, and a plurality of coolingopenings 508 that extend from face 506 to trailing edge 500. Coolingopenings 508 are oriented to facilitate channeling cooling air 526towards shroud segment leading edge 133 and to facilitate purging gap182 of migrating hot gases that have migrated into gap 182.

FIG. 4 is a schematic plan view of turbine nozzle outer band 183, outerband aft flange 504, gap 182, and shroud assembly leading edge 133. Morespecifically, as shown in FIG. 4, aft flange cooling openings 508 extendobliquely through aft flange 504 and shroud segment leading edgeopenings 181 extend obliquely through shroud segment leading edge 133.In the exemplary embodiment, each nozzle segment 520 includes at leastone airfoil vane 510 that includes a first sidewall 512 and a secondsidewall 514. In the exemplary embodiment, first sidewall 512 is convexand defines a suction side of each airfoil vane 510, and second sidewall514 is concave and defines a pressure side of each airfoil vane 510.Sidewalls 512 and 514 are joined together at a leading edge 516 and atan axially-spaced trailing edge 518 of each airfoil vane 510. Eachairfoil trailing edge 518 is spaced chordwise and downstream from eachrespective airfoil leading edge 516. First and second sidewalls 512 and514, respectively, extend longitudinally, or radially outwardly, in spanfrom a radially inner band (not shown) to radially outer band 183.

Each vane 510 has airfoil contours (not shown) from leading edge 516toward trailing edge 518. As the hot combustion gases flow aroundairfoil vanes 510, the gases along sidewall 512 accelerate and create alower static pressure and the gases along sidewall 514 decelerate andcreate a higher static pressure. During engine operation, hot combustiongases are channeled between vanes 510 and bands 183 and form a pair ofpassage vortices from sidewall 514 toward sidewall 512 on the surfacesof the inner band (not shown) and outer band 183. The passage vorticesbring the hotter combustion gases from the mid-span core flow toward theinner band (not shown) and outer band 183. There are periodic pressurevariations along the circumferential direction at the turbine nozzleouter band trailing edge 500. The combination of the passage vortex andthe circumferential pressure variation at shroud leading edge 133 resultin circumferentially periodic local hot spots 550 developing on shroudinner surface 138. Over time, such hot spots 550 may reduce an overallperformance of the engine assembly and/or reduce a durability of theengine.

Cooling openings 508 are obliquely-oriented in outer band aft flange 504relative to trailing edge 500 such that cooling flow discharged fromopenings 508 is discharged at a discharge angle α generally in thedirection of rotation indicated by arrow A. In the exemplary embodiment,discharge angle α is oblique and as such is not parallel to a flow ofcombustor gases through turbines 112 and 114. More specifically, allcooling openings 508 are obliquely-oriented at the same discharge angleα. Alternatively, any of openings 508 may be uniformlyobliquely-oriented at any discharge angle α that enables coolingopenings 508 to function as described herein. In the exemplaryembodiment, openings 508 are spaced circumferentially equidistantlyacross outer band trailing edge 500. Moreover, in the exemplaryembodiment, openings 508 are all sized and obliquely-orientedidentically. It should be appreciated that although cooling openings 508illustrated in the exemplary embodiment are sized identically and areuniformly spaced across trailing edge 500 of turbine nozzle outer band183, in alternative embodiments, cooling openings 508 may have any size,shape, or orientation that enables cooling openings 508 to function asdescribed herein.

In the exemplary embodiment, cooling openings 181 extend through shroudassembly leading edge 133 and are obliquely-oriented to dischargecooling fluid at a discharge angle β measured with respect to acenterline 555 of gap 182. In the exemplary embodiment, discharge angleβ is oblique and as such flow discharged from openings 181 is notparallel to a flow of combustion gases through turbines 112 and 114.More specifically, in the exemplary embodiment, cooling openings 181 areuniformly obliquely-oriented at discharge angle β in the direction ofarrow A. Alternatively, any of openings 181 may be uniformlyobliquely-oriented at any discharge angle β that enables coolingopenings 181 to function as described herein. In the exemplaryembodiment, openings 181 are spaced circumferentially equidistantlyacross shroud assembly leading edge 133. Moreover, in the exemplaryembodiment, openings 181 are all sized and oriented identically. Itshould be appreciated that although cooling openings 181 illustrated inthe exemplary embodiment are sized identically and are uniformly spacedacross the leading edge 133 of shroud assembly 110, in alternativeembodiments, cooling openings 181 may have any size, shape, ororientation that enables cooling openings 181 to function as describedherein.

In the exemplary embodiment, cooling openings 508 are each substantiallyaligned with a respective one of cooling openings 181 located across gap182. It should be appreciated that although cooling openings 508 aresubstantially aligned with respective cooling openings 181 in theexemplary embodiment, in other embodiments, cooling openings 508 are notrequired to align with respective cooling openings 181, and instead maybe offset by any distance from respective cooling openings 181 thatenables cooling openings 508 and 181 to function as described herein.Additionally, in the exemplary embodiment, discharge angles α and β havethe same magnitude. It should be appreciated that although angles α andβ are described as having the same magnitude in the exemplaryembodiment, in other embodiments, cooling openings 508 and 181 may beoriented at different angles α and β, respectively.

During operation, the oblique orientation of cooling openings 508 and181 imparts a clockwise, or tangential, velocity component to airchanneled through cooling openings 508 and 181. As a result, coolingflow energy is facilitated to be enhanced because little energy is lostin turning the cooling air through misaligned cooling openings. Theclockwise momentum of the air facilitates balancing the pressuredistribution inside gap 182, such that hot gas ingestion into gap 182 isreduced.

Moreover, the oblique orientation and location of cooling openings 508and 181 about turbine nozzle assembly 520 facilitates reducing hot gasingestion into gap 182 and facilitates improving film cooling acrossshroud inner surface 138 downstream from leading edge 133. Theorientation and location of cooling openings 508 and 181 facilitatesincreasing the length of cooling openings 508 and 181, thus increasingthe convective cooling ability of openings 508 and 181 within each outerband 183 and shroud assembly 110.

FIG. 5 is a schematic plan view of turbine nozzle outer band 183, outeraft flange 504, gap 182 and shroud assembly leading edge 133,illustrating an alternative embodiment of cooling openings 181 and 508.FIG. 6 is a schematic plan view of turbine nozzle outer band 183, outeraft flange 504, gap 182 and shroud assembly leading edge 133,illustrating another alternative embodiment of cooling openings 181 and508. FIG. 7 is a schematic plan view of turbine nozzle outer band 183,outer aft flange 504, gap 182 and shroud assembly leading edge 133,illustrating yet another alternative embodiment of cooling openings 181and 508. The engine assembly shown in FIGS. 5, 6, and 7 is the sameengine assembly shown in FIG. 4, with the exception of a few componentchanges, described in more detail below. As such, components illustratedin FIGS. 5, 6, and 7 that are identical to components illustrated inFIG. 4, are identified in FIGS. 5, 6, and 7 using the same referencenumerals used in FIG. 4.

With respect to FIG. 5, cooling openings 508 extend obliquely throughouter band aft flange 504 and two different groups 528 and 530 ofcooling openings 181 extend obliquely through shroud assembly leadingedge 133. During operation, hot spots 550 may develop on shroud innersurface 138 that may require enhanced local cooling. Consequently, inthis alternative embodiment, at least one group 528 of larger diametercooling openings 181 is positioned upstream from each corresponding hotspot 550. Openings 181 in group 528 facilitate preferential cooling bychanneling cooling fluid towards each corresponding hot spot 550. Itshould be appreciated that hot spot 550 locations may vary betweenengines and the locations of groups 528 will vary accordingly. In theexemplary embodiment, a group 530 of small diameter cooling openings 181is positioned between circumferentially-spaced groups 528 of largerdiameter cooling openings 181. More specifically, the relative locationsof groups 530 correspond to areas of shroud inner surface 138 that aredownstream from leading edge 133 and that are subjected to relativelylower operating temperatures in comparison to hot spots 550. Thus, inthe exemplary embodiment of FIG. 5, cooling opening groups 528 and 530may be positioned to enhance a flow of cooling fluid provided to hotterregions of inner surface 138 while minimizing a flow of cooling fluidprovided to cooler regions of inner surface 138, thereby facilitatingreducing hot spot development on shroud inner surface 138. It should beappreciated that the term “fluid” as used herein includes any coolingmaterial or medium that flows, including but not limited to, gas, airand liquids, as described herein.

With respect to the embodiment illustrated in FIG. 6, at least one group532 of larger diameter cooling openings 508 is positioned upstream fromeach corresponding hot spot 550, to facilitate preferential cooling ofeach corresponding hot spot 550. It should be understood that in thisembodiment, shroud assembly leading edge 133 does not include coolingopenings 181. It should also be appreciated that hot spot 550 locationsmay vary between engines and as such, locations of groups 532 will varyaccordingly. In the exemplary embodiment, a group 534 of small diametercooling openings 508 is positioned between circumferentially-spacedgroups 532 of larger diameter cooling openings 508. More specifically,the relative locations of groups 534 correspond to areas of shroud innersurface 138 that are downstream from leading edge 133 and that aresubjected to relatively lower operating temperatures in comparison tohot spots 550. Thus, in the exemplary embodiment of FIG. 6, coolingopening groups 532 and 534 may be positioned to enhance a flow ofcooling fluid provided to hotter regions of inner surface 138 whileminimizing a flow of cooling fluid provided to cooler regions of innersurface 138, thereby facilitating reducing hot spot development onshroud inner surface 138.

With respect to the embodiment illustrated in FIG. 7, the coolingopening patterns shown in FIGS. 5 and 6 are essentially combined. Morespecifically, groups 528 and 532 of larger diameter cooling openings 181and 508, respectively, are positioned upstream from each correspondinghot spot 550, to facilitate preferential cooling by channeling coolingfluid towards each corresponding hot spot 550. Groups 528 and 532 arepositioned to cooperate with each other across opposite sides of gap182. Likewise, groups 530 and 534 are positioned to cooperate with eachother across opposite sides of gap 182. It should be appreciated thathot spot 550 locations may vary between engines and the locations ofgroups 528 and 532 will vary accordingly. In the exemplary embodiment,groups 530 and 534 of small diameter cooling openings 181 and 508,respectively, are positioned between circumferentially-spaced groups 528and 532, respectively, of large diameter cooling openings 181 and 508,respectively. Moreover, the relative locations of groups 530 and 534correspond to areas of shroud inner surface 138 that are downstream fromleading edge 133 and that are subjected to relatively lower temperaturesin comparison to hot spots 550. Thus, in the exemplary embodiment ofFIG. 7, groups 528, 530, 532 and 534 may be positioned to enhance a flowof cooling fluid provided to hotter regions of inner surface 138 whileminimizing a flow of cooling fluid provided to cooler regions of innersurface 138, thereby reducing hot spot development on shroud innersurface 138.

It should be appreciated that large diameter cooling openings 181 and508 of groups 528 and 532, respectively, may be any diameter thatenables groups 528 and 532 to function as described herein. Moreover, itshould be appreciated that small diameter cooling openings 181 and 508of groups 530 and 534, respectively, may be any diameter that enablesgroups 530 and 534 to function as described herein.

It should be appreciated that although the aforementioned exemplaryembodiments describe patterns of cooling openings 181 and 508 withingroups 528 and 530, and 532 and 534, respectively, other embodiments mayuse different patterns of cooling openings 181 and 508 within respectivegroups 528 and 530, and 532 and 534. More specifically, otherembodiments may facilitate preferential cooling of hot spots 550 byadjusting, in any manner, the aggregate cross-sectional area availablefor channeling cooling fluid to hot spots 550 within each group 528,530, 532 and 534. For example, other embodiments may use the samediameter for cooling openings 181 and 508 within respective groups 528and 530, and 532 and 534, and at the same time increase or decrease thedensity of cooling openings 181 and 508 within respective groups 528 and530, and 532 and 534. Consequently, to increase cooling flow through anyof groups 528, 530, 532 and 534, the density of respective coolingopenings 181 and 508 is increased. To decrease cooling flow through anyone of groups 528, 530, 532 and 534, the density of respective coolingopenings 181 and 508 is decreased. Thus, by adjusting the density ofcooling openings 181 and 508 within respective groups 528 and 530, and532 and 534, the aggregate cross-sectional area is also adjusted andpreferential cooling of hot spots 550 is facilitated.

In another example of adjusting the aggregate cross-sectional areaprovided for channeling cooling fluid, groups 528 and 532 may includeany number of large diameter cooling openings 181 and 508, respectively,at any cooling opening spacing, that enables groups 528 and 532 tofunction as described herein. Likewise, groups 530 and 534 may includeany number of small diameter cooling openings 181 and 508, respectively,at any cooling opening spacing, that enables groups 530 and 534 tofunction as described herein.

In yet another example of adjusting the aggregate cross-sectional areaprovided for channeling cooling fluid, groups 528 and 532 may include anincreased density of large diameter cooling holes 181 and 508,respectively, that enables groups 528 and 532 to function as describedherein. Similarly, the aggregate cross-sectional area provided forchanneling cooling fluid, groups 530 and 534 may include a decreaseddensity of small diameter cooling holes 181 and 508, respectively, thatenables groups 528 and 532 to function as described herein.

It should be appreciated that although the exemplary embodimentsdescribe cooling openings 181 and 508 as having circular cross-sections,other embodiments may use any cross-sectional area for cooling openings181 and 508 that enables respective groups 528 and 530, and 532 and 534,to function as described herein. Such cross-sectional areas include, butare not limited to, oval, square and rectangle. It should be appreciatedthat cooling openings 181 and 508 of respective groups 528 and 532 arepositioned upstream of hot spot 550. However, due to the oblique gasflow exiting from vanes 510, openings 181 and 508 are not necessarilyaxially aligned with the engine.

The above-described turbine nozzle segments and shroud segments includea plurality of obliquely-oriented cooling openings extending along anaft flange of the turbine nozzle outer band and along a leading edge ofthe turbine shroud assembly. More specifically, cooling openings extendthrough the aft flange of the outer band and through the leading edge ofthe turbine shroud assembly, such that groups of cooling openings havinga greater aggregate cross-sectional area are positioned upstream oflocal hot spots and groups of cooling openings having a smalleraggregate cross-sectional area are positioned between the groups havinggreater aggregate cross-sectional area. As a result, cooling fluid isdirected into a gap defined between the aft flange and leading edge tofacilitate reducing hot gas ingestion into the gap, and to facilitateenhanced local cooling by distributing cooling fluid to hot spots on theshroud inner surface. Accordingly, the turbine nozzle segments andshroud segments are operable at a reduced operating temperature, thusfacilitating extending the durability and useful life of the turbinenozzle segments and shroud segments, and reduces the operating cost ofthe engine.

Exemplary embodiments of turbine nozzle segments and shroud segments aredescribed above in detail. The segments are not limited to the specificembodiments described herein, but rather, components of each segment maybe utilized independently and separately from other components describedherein.

While the invention has been described in terms of various specificembodiments, those skilled in the art will recognize that the inventioncan be practiced with modification within the spirit and scope of theclaims.

1. A method of assembling a gas turbine engine, said method comprising:coupling at least one turbine nozzle segment within the gas turbineengine, wherein the at least one turbine nozzle segment includes atleast one airfoil vane extending between an inner band and an outer bandthat includes an aft flange and a radial inner surface; coupling atleast one turbine shroud segment downstream from the at least oneturbine nozzle segment, wherein the at least one turbine shroud segmentincludes a leading edge and a radial inner surface; coupling a coolingfluid source in flow communication with the at least one turbine nozzlesegment such that cooling fluid channeled to each turbine nozzle outerband aft flange is directed at an oblique discharge angle towards theleading edge of the at least one turbine shroud segment; and channelingthe cooling fluid through at least a first group of cooling openingshaving a first aggregate cross-sectional area and a second group ofcooling openings having a second aggregate cross-sectional area, thefirst aggregate cross-sectional area is larger than the secondcross-sectional area, to facilitate preferential cooling of the turbineshroud.
 2. A method in accordance with claim 1 wherein said coupling acooling fluid source in flow communication with the at least one turbinenozzle segment further comprises coupling the cooling fluid source tothe turbine nozzle segment such that cooling fluid is channeled into agap defined between the aft flange and the leading edge.
 3. A method inaccordance with claim 1 wherein channeling the cooling fluid furthercomprises channeling cooling fluid through the first group of openingsto facilitate enhanced cooling fluid flow to hotter portions of theshroud segment inner surface.
 4. A method in accordance with claim 1wherein channeling the cooling fluid further comprises channelingcooling fluid through the second group of openings to facilitatereducing cooling fluid flow to cooler portions of the shroud segmentinner surface.
 5. A method in accordance with claim 1 further comprisingforming the first group of cooling openings from a plurality of coolingopenings having a first diameter and forming the second group of coolingopenings from a plurality of cooling openings having a second diameter,the first diameter is larger than the second diameter.
 6. A method inaccordance with claim 1 further comprising forming the first group ofcooling openings from an a first density of cooling openings and formingthe second group of cooling openings from a second density of coolingopenings, the first density of cooling openings is greater than thesecond density of cooling openings.
 7. An engine assembly comprising: anozzle assembly comprising: an inner band; an outer band comprising anaft flange and a radial inner surface, said aft flange comprising afirst group of cooling openings having first aggregate cross-sectionalarea and a second group of cooling openings having a second aggregatecross-sectional area, said first aggregate cross-sectional area islarger than said second aggregate cross-sectional area, the coolingopenings configured to direct cooling fluid therefrom at an obliquedischarge angle; and at least one airfoil vane extending between saidinner band and said outer band.
 8. An engine assembly in accordance withclaim 7 further comprising a shroud assembly, said shroud assemblycomprising a shroud inner surface and a shroud assembly leading edge,said shroud assembly leading edge comprising a first shroud group ofcooling openings having a third aggregate cross-sectional area and asecond shroud group of cooling openings having a fourth aggregatecross-sectional area, said third aggregate cross-sectional area islarger than said fourth aggregate cross-sectional area, said coolingopenings configured to discharge cooling fluid therefrom at an obliqueangle with respect to a centerline of a gap defined between said outerband aft flange and said shroud assembly leading edge.
 9. An engineassembly in accordance with claim 8 wherein said first group of coolingopenings and said first shroud group of cooling openings include aplurality of cooling openings having a first diameter.
 10. An engineassembly in accordance with claim 8 wherein said second group of coolingopenings and said second shroud group of cooling openings include aplurality of cooling openings having a second diameter, said firstdiameter is larger than said second diameter.
 11. An engine assembly inaccordance with claim 8 wherein said first group of openings and saidfirst shroud group of openings include a plurality of densely orientedcooling holes.
 12. An engine assembly in accordance with claim 8 whereinsaid second group of openings and said second shroud group of openingsinclude a plurality of cooling holes less densely oriented than saidfirst group of openings and said first shroud group of openings.
 13. Anengine assembly in accordance with claim 8 wherein said first group ofcooling openings and said second group of cooling openings facilitatereducing hot gas ingestion into a gap defined between said outer bandaft flange and said shroud assembly leading edge.
 14. An engine assemblyin accordance with claim 8 wherein said first group of cooling openingsand said first shroud group of cooling openings facilitate film coolingof said shroud inner surface.
 15. A gas turbine engine comprising anozzle assembly comprising an inner band, an outer band, and at leastone airfoil vane extending between said inner band and said outer band,said outer band comprising an aft flange and a radial inner surface,said aft flange comprising a first group of cooling openings having afirst aggregate cross-sectional area and a second group of coolingopenings having a second aggregate cross-sectional area, said firstaggregate cross-sectional area is larger than said second aggregatecross-sectional area, said cooling openings configured to direct coolingfluid therefrom at an oblique discharge angle.
 16. A gas turbine enginein accordance with claim 15 further comprising a shroud assembly, saidshroud assembly comprising a shroud inner surface and a shroud assemblyleading edge, said shroud assembly leading edge comprising a firstshroud group of cooling openings having a third aggregatecross-sectional area and a second shroud group of cooling openingshaving a fourth aggregate cross-sectional area, said third aggregatecross-sectional area is larger than said fourth aggregatecross-sectional area, said cooling openings configured to dischargecooling fluid therefrom at an oblique angle with respect to a centerlineof a gap defined between said outer band aft flange and said shroudassembly leading edge.
 17. A gas turbine engine in accordance with claim16 wherein said first group of cooling openings and said first shroudgroup of cooling openings include a plurality of cooling openings havinga first diameter.
 18. A gas turbine engine in accordance with claim 17wherein said second group of cooling openings and said second shroudgroup of cooling openings include a plurality of cooling openings havinga second diameter, said first diameter is larger than said seconddiameter.
 19. A gas turbine engine in accordance with claim 16 whereinsaid first group of cooling openings facilitates reducing hot gasingestion into a gap defined between said outer band aft flange and saidshroud assembly leading edge.
 20. A gas turbine engine in accordancewith claim 16 wherein said first group of cooling openings and saidfirst shroud group of cooling openings facilitate film cooling of saidshroud inner surface.